An example of an attitude and heading measurement system is described in document FR 2 614 694, which corresponds to European Patent No. 0292339.
Document FR 2 614 694 describes a unit with three detection assemblies respectively delivering gyro components, acceleration components, and magnetometer components in a three-axis system tied to the body of an aerodyne. On the basis of the gyro information, computation means generate a projection matrix for projecting these components into the local navigation three-axis system.
Other means enable the magnetometer components to be projected into the local navigation system using said matrix, and then to determine the heading error in the longitudinal-and-transverse axis system.
Furthermore, means generate corrections for the gyro measurements in the axis system tied to the aerodyne body on the basis of said attitude and heading errors.
In the attitude and heading measurement systems according to document FR 2 614 694, the estimated gyro bias, i.e. the difference between a true value of a gyro variable and the measured value, is frozen whenever the aircraft is turning, e.g. at more than 0.6 degrees per second (°/s).
The purpose of freezing the estimated gyro bias is to avoid instability that occurs in the attitude filter as a result of turning, and that would lead to divergence of the attitude estimates.
The drawback of that approach is that any variations in gyro measurement bias that take place after the bias has been frozen are no longer compensated. In general, this has little consequence when the gyros are of tactical grade, as applies to fiber optical gyros (FOGs) that present precision that is conventionally of the order of one degree per hour (1°/h).
In contrast, when the gyros of an attitude unit are of lower performance, that can lead to said attitude unit producing significant errors in its estimates of attitudes and of heading.
This applies for example when gyro bias is of the order of several tens of degrees per hour, as it is with vibrating gyros. Vibrating gyros in the form of micro-electromechanical systems (MEMSs) are presently available. Such MEMS gyros can be manufactured using methods that are close to those used for manufacturing integrated circuits, and are low in cost.
One approach to this problem provides for hybridizing data from a positioning system such as the global positioning system (GPS) or the Galileo positioning system, or the like, and data coming from an inertial measurement unit (IMU) in order to constitute an inertial navigation system (INS).
Document FR 2 898 196 describes such an approach, as does the document “Reconstitution de l'état d'un micro drone par fusion de données” [Reconstructing the state of a micro-drone by data merging] by Joan Solà Ortega, available at: http://homepages.laas.fr/jsola/JoanSola/objectes/a ltres/DEA/RapportDEA.doc
According to those documents, a global filter is proposed (attitudes/heading, speed, position/altitude). The attitude, heading, and vertical filters are not separable into sub-filters, which prevents them from being resolved, as is essential for proving that they have the stability required, e.g. for certification that is “safety critical”. Furthermore, the three Euler angles (cf. infra) are handled in the complex form of a quaternion of variable Q and dimension four (4). Finally, the biases of the sensors, in particular of the gyros, are expressed in a body frame of reference, thereby complicating the formulation thereof.
Furthermore, it is presently not possible for an attitude unit to include gyros presenting estimated error or bias that is of the order of several tens of degrees per hour. A fortiori, this is completely impossible if it is desired to obtain safety-critical certification. Nevertheless, low-cost vibrating gyros exist in MEMS form. With present techniques, such MEMS gyros present biases that are too great for them to be included in an attitude unit.